1. Field of the Invention
The present invention relates to vehicle control and more particularly to attitude control apparatus for spacecraft. Yet more specifically, the present invention relates to a novel arrangement for reaction wheels (RW""s), momentum wheels (MW""s) and/or control moment gyros (CMG""s) used to position a space vehicle and control its attitude.
2. Description of the Prior Art
RW""s, MW""s and CMG""s have long been used to control the attitude of space craft. They were designed to replace reaction jets for cyclic type maneuvers and provided improved control, longer spacecraft life and reduced fuel requirements. In the prior art, a number, at least three and usually more for redundancy and fail safe operation, have been mounted at various locations about the spacecraft as dictated by volume, structure and thermal considerations and where sufficient space was available. The components, being remote from one another, have a number of disadvantages. For example, several boxes of electronics are required for each component and this introduces greater weight, volume, cost and heat generation. With the prior art each CMG, MW or RW has to be installed separately and many of the functions have to be calibrated and tested by the spacecraft integrator. The prior CMG art requires a gimbal rate sensor or tachometer which when combined with other data can be used to derive an approximation of the torque being delivered to the spacecraft. The sensor its self induces higher frequency signals which results in added and undesirable vibrations being transmitted to the spacecraft. Finally, the prior CMG art uses ring like mounting structures which are bolted to a plate structure in the spacecraft. This primarily two-dimensional structure is structurally inefficient. Each unit delivers torque to the spacecraft as its single function, but the net effect cannot be measured directly and efficiently. As will be shown, the present invention is more of a three-dimensional system having somewhat equal dimensions in all three spatial axes.
The present invention overcomes the problems of the prior art by providing an integrated single unitary structure containing multiple spinning bodies, the electronics to control them and the intelligence to interface with the spacecraft on the sub-system level. Repackaging the system into a single integrated unit allows the manufacturer to design a single and more efficient structure to house the inner gimbal assemblies of the RW""s, MW""s or CMG""s. With CMG""s the outer gimbal base ring as a separate part, can be eliminated and the lower number of electronic boxes can be reduced resulting in reduced weight and the number of connections to the vehicle further reduces weight and improves heat conduction transfer to the outer surfaces of the vehicle. Further more, the unit manufacturer now has to deal with a single momentum control unit and, on delivery, the unit will be able to be plugged in to reduce the contractor cost and improve the spacecraft manufacturing schedule. The total volume occupied by the unitary structures will be smaller than that of separate mounted units. The arrangement and other design changes enabled by the grouping together converts the result into a higher level system offering many advantages including, better performance, the ability to measure the performance more accurately and more directly, lower weight, lower power, a smaller package, and lower cost. Using common circuits within the system reduces the number of electronic components. The number and resulting weight of cables and connector is also be reduced. Grouping them and adding a single six axes interfacing kinematics force measuring system, such as is proposed herein, enables a much more effective servo control system to regulate the resulting net torque and consequential motion of the spacecraft. This further enables the addition of a single isolation system as an integral part of the force measuring system to reduce or filter undesirable high frequency vibratory forces that are otherwise transmitted to the spacecraft. The present invention not only avoids this problem it also diminishes the effects of other sources of high frequency vibrations such as that caused by the bearings and rotor unbalance. The present invention allows the manufacturer to complete this process much more professionally before it leaves the factory and in parallel with the building of the spacecraft. By operating on a higher level set of requirements the manufacturer can relax sub-level requirement lowering his manufacturing cost. Installation into the spacecraft is much simpler. The result is lower spacecraft integrator cost and significantly reduced spacecraft manufacturing schedules which further reduces cost.